This invention relates to a spacecraft propulsion system and particularly to a propulsion system which efficiently integrates the apogee boost and reaction control propulsion system components.
Spacecraft such as unmanned satellites which must be boosted into high earth orbits such as geostationary orbits are typically injected into a low earth orbit, for example, by an unmanned launch vehicle or by a manned space shuttle orbiter. Once in the low altitude orbit, many satellites must be boosted by their own propulsion system to reach the desired final orbital position. In one type of prior art propulsion system, the satellite is boosted by first firing a solid rocket perigee kick motor which is jettisoned after it is exhausted and final orbit is achieved by firing a liquid fuel apogee kick motor (AKM). New generation satellites employ a single liquid fuel AKM to perform the orbit transfer function. Satellites boosted into high earth orbit are also equipped with a number of relatively small thrust motors which comprise the reaction control system (RCS) which is used to make fine changes in spacecraft orbit and position and for station keeping. In the design of satellites and their propulsion systems, designers are constantly striving to improve the efficiency, performance, reliability, and operational lifespan of such propulsion systems.
One type of present day satellite propulsion system integrates the fuel supply systems of the RCS and AKM. Integration is desirable to enable the RCS to take advantage of residual fuel not expended by the AKM which would otherwise be unusable if the systems were separated. Both of these propulsion systems operate by controlled mixing of a liquid fuel and oxidizer (hereinafter collectively referred to as propellant) which produces a hypergolic reaction (combustion upon contact). Fluid tanks filled with liquid fuel and oxidizer have a gas space which are initially pressurized to a high level. During AKM firing, the fuel and oxidizer tanks pressure is regulated to maintain this high pressure level using an external source of gas pressure (typically helium). When AKM firing is completed, pressure regulation of the fuel tanks is ended and the remaining propellant in the tanks is used by the RCS thrusters operating in what is known as a "blowdown" mode. As the RCS thrusters periodically consume fuel within the tanks, propellant pressure is gradually reduced to a point where efficient and reliable operation of the RCS thrusters is no longer possible. Therefore, at a given pressure level, the RCS must be deactivated even though a supply of fuel and oxidizer remains. Consequently, a greater mass of fuel and oxidizer must be carried onboard the spacecraft to insure that sufficient usable fuel and oxidizer are available for the desired RCS functions.
The present day integrated spacecraft propulsion systems of the type described above have a number of disadvantages. Since the fuel and oxidizer tanks must be designed to withstand high internal pressures, they must be very strong, thick-walled tanks having considerable mass. Furthermore, the tank pressurization system must be fairly complex in order to provide the desired level of pressurization and pressure control. The existence of residual fuel and oxidizer within the tanks decreases the total stage efficiency of the spacecraft (defined as the mass of the initial usable propellant divided by the total stage mass). Finally, thrusters which must be designed to operate over a range of supply pressures require design compromises which produce less than optimal performance. The above considerations impose fundamental limits in propulsion system efficiency and performance achievable with current integrated liquid fuel satellite propulsion systems.
Designers of spacecraft propulsion systems are now turning their attention toward designing pump-fed propulsion systems. Pump-fed systems employ a fuel-pressurizing pump which receives liquid fuel or oxidizer from the storage vessels and supplies it under pressure to the thrusters. For pump-fed systems, the liquid fuel-containing vessels need be pressurized to only a fairly low level since the only requirement is that fuel supplied to the pump must not cavitate during pumping operation. The advantages of pump-fed systems include higher performance and greater stage efficiencies. The higher stage efficiencies result from the use of lighter, thin-walled liquid tanks and a smaller, less complex tank pressurization system. Further, the pump-fed systems permit usage of nearly all of the liquid within the vessels and provide nearly uniform propellant pressures. Although the above-mentioned advantages are provided for the AKM, problems are encountered in designing the relatively small RCS thrusters to operate with the pump-fed AKM. The small RCS thrusters cannot operate from the same pumps as the AKM, as it is impractical to start and stop these relatively large pumping capacity units for each RCS maneuver. Small pumps for the RCS are not currently available and the low pressure within the propellant tanks is insufficient to operate conventional RCS thrusters. The development, testing and qualification of new RCS pumps or low pressure RCS thrusters would be an extremely expensive and time-consuming process. Therefore, it is currently impractical to integrate a pressure-fed RCS into a pump-fed AKM without modifications to the propulsion system that defeat some of the advantages inherent with the pump-fed system.
During the course of development of this invention by the inventors, a number of alternate design solutions were considered. One proposed improved propulsion system would employ a pump-fed AKM which would achieve benefits in terms of AKM performance (due to constant propellant supply pressure) and increase stage efficiency since thin-walled lightweight propellant tanks could be used. Once the AKM function is complete, the system would use a special low pressure RCS thruster which would operate in the same fashion as a conventional pressure-fed system; that is, the thruster would operate in a straight blowndown mode until its lowest operating pressure level is reached. The advantages of such a design proposal are that the system integrates the AKM and RCS propellant systems, it is no more complex than the present systems, and the stage efficiency is improved compared to current systems. This design solution, however, has several disadvantages. First, a new low pressure RCS thruster would have to be developed which would be a major and expensive undertaking. Second, thermal requirements for the propellant feed system are extremely stringent. Thruster inlet pressures of 50 psi gives RCS thruster chamber pressures of about 20 psi. The vapor pressure of one commonly used oxidizer at 100 degrees F. is 33 psi, and at 82 degrees, F. is 20 psi. If the chamber pressure equalled or approached the vapor pressure, the liquid would undergo a phase change which would interrupt operation of the thruster. Therefore, the need for precise thermal control would be necessary in order to prevent fuel vaporization within the thruster chambers. Finally, lower performance of a low pressure RCS thruster would be anticipated as compared within conventional high pressure RCS thrusters. This lower performance of the low pressure RCS thruster would increase the propellant requirements for altitude control with a compounding effect on the propellant mass necessary for AKM operation.
Another proposed design solution was to provide a non-integrated system which employs a pump-fed AKM and a separate RCS propulsion system which operates in a straight blowdown mode. Since the propellant tanks of the blowdown RCS propulsion system are considerably smaller than those needed to supply the AKM, the mass disadvantage of high pressure tankage is minimized. The pressure blowdown range of the RCS system could be from about 350 to 100 psi. The advantages of such a blowdown system is its simplicity, since no new components are needed. The major disadvantages of such a design approach are that the residuals in the AKM tanks are unusable in the RCS (because the systems are separate) and that the blowdown range of the RCS thrusters is large, giving a lower performance over the life of the spacecraft.
Still another design proposal was to incorporate an accumulator that would use either the AKM fuel pump or a smaller auxiliary pump to refill small high-pressure tanks. These tanks could be blown down normally and refilled using the pump which could be powered by spacecraft batteries. The advantages of such a system are that the primary tanks and helium system are optimized for a pump fed system. The disadvantage of this system is that the auxiliary pumps must be operated several times during the spacecraft life, which is typically about ten years. A pump design that can reliably operate over a ten-year mission has not yet been developed or qualified. If AKM pumps were used, the system would be inefficient since they have a capacity much larger than required for filling small RCS accumulators. Finally, system redundancy requires multiple pumps and highly complex manifolding and electrical systems. Accordingly, this design proposal was also determined to be unacceptable.
In view of the foregoing, there is a need to provide a highly efficient, integrated RCS/AKM propulsion system which provides the advantages of a pump-fed AKM without imposing severe limitations to the RCS function or require design changes which defeat the advantages of a pump-fed AKM.